1. Field of the Invention
The invention relates to high-pressure turbine nozzles of gas-turbine engines, in particular aircraft gas-turbine engines.
2. Description of the Related Art
It is known that increasing the temperature at the intake of the turbine(s) of a gas-turbine engine allows optimization of the performance of the gas-turbine engine. Increasing the temperature provides a gain in specific fuel consumption, increasing the aircraft's range or decreasing the amount of fuel that must be carried along. This temperature increase also increases the thrust of the gas-turbine engine. Present-day engines can operate with a temperature of 1,577.degree. C., whereas the gas-turbine engines of 1950 (for instance the ATAR design) were limited to 930.degree. C.
A cooling system for the nozzle vanes and the turbine wheel is required for operating at such temperatures. To this end, a systematic circulation of cooling air is established inside the vane and perforations are formed in the vane wall which are suitably configured to create a vane protective film. Cooling is implemented by two substantial methods, namely internal convection and the protective film.
FIGS. 1 and 2 show the design presently used on the CF6-80 and GE 90 engines.
The vane comprises a hollow airfoil 1 inserted between a hollow outer deck and a hollow inner deck 9. The airfoil 1 contains a liner 2 defining a continuous peripheral cavity 3 between a wall 4 of the airfoil 1 and the outside of the liner 2. Spacers affixed to the wall 4 or the liner 2 keep the liner 2 and the wall 4 apart. The liner 2 comprises a multi-perforated skirt 6 and an inner wall 7 spaced a distance away from a wall 8 separating the airfoil 1 from the inner deck 9. The wall 8 is fitted with an orifice 10. An airflow 11 coming from a compressed-air source, generally the compressor of the gas-turbine engine, passes through the outer deck, enters the inside of the liner 2 and exhausts through the multi-perforation of the skirt 6 forming air jets 11' in the peripheral cavity 3 to impact-cool the wail 4 of the airfoil 1. Next, this airflow 11 passes inside and across the inner deck 9, which is cooled thereby, to finally exhaust as cooling airflow 19 through orifices 12 located at the upstream side of the inner deck 9. The pressure P1 in an upstream enclosure 13 is larger than the pressure in a downstream enclosure 14. These enclosures are bounded by the inner deck 9 and an upstream turbine disk 15 and a downstream turbine disk 16, respectively. An airflow will be set up through a labyrinth 18 separating the enclosures 13 and 14 which determines the magnitude of an airflow 17. This airflow 17 is expelled into the path of hot gases downstream of the nozzle to cool the periphery of the downstream disk 16. The remainder of the cooling airflow 19 coming from the orifices 12 is expelled into the path of hot gases upstream of the nozzle to cool the periphery of the upstream turbine disk 15. In this manner, these airflows 17 and 19 cool the high-pressure turbine disks. By this simple technical configuration, the total cooling airflow 11 injected inside the liner 2 initially acts to cool the airfoil 1.
However, the air injected inside the airfoil 1 must descend again to the airfoil base to be exhausted through the inner deck 9. As a result, the flow for impact-cooling is sheared, impeding mathematical modeling. Because the air cooling the inner deck 9 and the turbine disks 15, 16 was already heated when cooling the wall 4 of the airfoil 1, the cooling of the inner deck 9 and the turbine disks 15, 16 is degraded.
If a crack 20 materializes at the trailing edge of the airfoil 1 in the manner shown in FIG. 2, at least a portion 21 of the flow of cooling air will pass through the crack 20. Since the pressure downstream of the vane is less than the pressure upstream, there is a danger of flow reversal inside the inner deck 9 which is crossed by a flow 22 of hot gases coming from the upstream part of the nozzle and passing through the upstream enclosure 13. In this case, the inner deck 9 will be heated and serious damage to, even destruction of, the vane may ensue. The flow underneath the inner deck 9 arises directly from the very hot gas of the gas flow path, entailing dangerous rotor heating and possibly destruction of the entire turbine if the inner deck 9 is not cooled properly.
FIGS. 3 and 4 show another embodiment of the cooling circuit which is similar to the embodiment of FIGS. 1 and 2. In this case, the inner wall 7 of the liner 2 comprises an outlet orifice 23 opposite an intake orifice 10 in the inner deck 9. Accordingly, the cooling airflow 11 enters the liner 2 and a portion 24 of this cooling air directly crosses the outlet orifice 23 to circulate in and cool the inner deck 9. As in the embodiment shown in FIGS. 1 and 2, the air circulating in the inner deck 9 is exhausted through the orifice 12 situated upstream of the inner deck 9, with one portion 17 crossing the labyrinth 18 while the other portion 19 is expelled into the path of hot gases upstream of the nozzle, thereby cooling the turbine disks 15 and 16.
In this second design, the air cooling the inner deck 9 and the disks 15, 16 is desirably fresher and the shearing flow of the impact-cooling air is less. On the other hand, the airfoil 1 is disadvantageously cooled by a lesser flow. Moreover, leaks 25, 26 may escape from the outlet orifice 23 into the peripheral cavity 3, complicating mathematical flow modeling.
If a crack 20 appears at the trailing edge of the airfoil 1 in the manner shown in FIG. 4, a portion 21 of the cooling air will escape into the hot-gas flow and, depending on engine operating conditions, the air's magnitude of flow in the inner deck 9 will be reduced or it's direction will be reversed. The inner deck 9 and the airfoil 1 also may be damaged or destroyed. Overheating in the enclosures 13 and 14 may dangerously raise the rotor temperature causing disk expansion and leading to contact between movable components and stationary rings.
According to the two embodiments of the state of art described above, the wall 4 of the airfoil 1 is furthermore fitted with gauged orifices to form a protective film around the airfoil.